IT202100029963A1 - GAS TURBINE ENGINE INCLUDING A ROTATING BLADE ASSEMBLY. - Google Patents

Description

GAS TURBINE ENGINE INCLUDING A ROTATING BLADE ASSEMBLY

TECHNICAL FIELD

[0001] This disclosure generally concerns a gas turbine engine and more? specifically a gas turbine engine rotating blade assembly.

STATE OF ART

[0002] Turbine engines, and in particular gas turbine engines, are rotary engines that extract energy from a flow of working air passing in series through a compressor section, where the working air is compressed, a combustor section, where fuel is? added to the working air and ignited, and a turbine section, where the burnt working air is expanded and work? taken from the working air to drive the compressor section along with other systems and provide thrust in an aircraft deployment. Compressor and turbine mounts include axially arranged pairs of rotating blades and stationary vanes. Can the gas turbine engine? be arranged as an engine center including at least one compressor section, one combustor section and one turbine section in axial flow arrangement and defining at least one rotating element or rotor and at least one stationary component or stator. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A complete and implementing disclosure of this description, including its modality? best, directed to a technician of the branch of ordinary competence,? presented in the specification, which refers to the attached figures, in which:

[0004] Figure 1 ? a schematic cross-sectional diagram of an aircraft turbine engine including a counter-rotating turbine section and a counter-rotating compressor section in accordance with various aspects described herein.

[0005] Figure 2 ? a cross-sectional view of the counter-rotating turbine section of the turbine engine of Figure 1 viewed from cut II of Figure 1, further illustrating a rotating blade assembly including a disk, a blade assembly and a retainer assembly.

[0006] Figure 3 ? an exploded perspective view of the rotary blade assembly of Figure 2, further illustrating the disk, the blade assembly and the retainer assembly.

[0007] Figure 4 ? a cross-sectional view of the rotating blade assembly viewed from cut IV-IV of Figure 3, further illustrating the retainer assembly.

[0008] Figure 5 ? a cross-sectional view of an exemplary rotating blade assembly of the gas turbine engine of Figure 1, further illustrating an exemplary disk and an exemplary blade assembly.

[0009] Figure 6 ? a radial view of the rotating blade assembly of Figure 5, further illustrating a tail of the exemplary disk received within the exemplary blade assembly.

[0010] Figure 7 ? a cross-sectional view of an exemplary rotor blade assembly of the gas turbine engine of Figure 1, further illustrating an exemplary disk and an exemplary retainer assembly including a retainer plate.

[0011] Figure 8 ? a cross-sectional view of an exemplary rotor blade assembly of the gas turbine engine of Figure 1, further illustrating an exemplary disk and an exemplary retainer assembly including a retainer plate.

[0012] Figure 9 ? a schematic axial view of an exemplary rotating blade assembly of the gas turbine engine of Figure 1, further including an exemplary disk and a set of retainers provided along a portion of the disk corresponding to each other blade assembly.

DETAILED DESCRIPTION

[0013] Aspects of this disclosure relate to a rotary blade assembly for a gas turbine engine including a drive shaft. The rotating blade assembly further comprising a disk operatively coupled to the drive shaft and comprising a seat having at least a portion of the first through hole and at least one blade assembly having a dovetail. A complex of holding elements can? secure at least one set of blades to the disc. Aspects of this disclosure are described in terms of a gas turbine engine, specifically a gas turbine engine including a counter-rotating section. In other words, a counter-rotating gas turbine engine. As used herein, the term “counter-rotating section,” or iterations thereof, may refer to a portion of the gas turbine engine comprising a series of axially adjacent, serially arranged rotating components (e.g., blades) that rotate in opposite circumferential directions. It is understood, however, that, although described in terms of the counter-rotating gas turbine engine, aspects of the disclosure described herein are not limited in this regard and may have applicability? general within any suitable gas turbine engine, a turboprop, turboshaft or turbofan engine having a power gear, in non-limiting examples. Will it be understood? Furthermore, however, that aspects of disclosure described herein are not so limited and may have applicability? general within other gas turbine engines. For example, disclosure can have applicability? for a complex of rotating blades in other engines or vehicles and can? be used to provide benefits in industrial, commercial and residential applications.

[0014] As used herein, the term ?anterior? or ?upstream? refers to a movement in one direction toward the gas turbine engine input or component that is relatively more? closer to the gas turbine engine inlet than another component. The term ?posterior? or ?downstream? used in conjunction with ?anterior? or ?upstream? refers to a direction towards the rear or exit of the gas turbine engine or which ? relatively more? closer to the gas turbine engine outlet than another component.

[0015] As used herein, ?a series? can? include any number of the elements respectively described, including a single element. In addition, the terms ?radial? or ?radially? as used herein, they refer to a quantity extending between a central longitudinal axis of the gas turbine engine and an external circumference of the engine.

[0016] All directional references (e.g. radial, axial, proximal, distal, superior, inferior, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, front, back, etc.) are used for identification purposes only to facilitate the reader's understanding of this disclosure and do not create limitations, particularly regarding the location, orientation or use of the disclosure. Link references (e.g. attached, coupled, connected, and joined) are to be interpreted broadly and may include intermediate members between a collection of elements and a reciprocal movement between elements, unless otherwise specified. Therefore, link references do not necessarily mean that two elements are connected directly and in a fixed relationship to each other. The exemplary drawings are for illustrative purposes only and the sizes, positions, order and reciprocal dimensions reflected in the attached drawings may vary.

[0017] Figure 1 ? a schematic cross-sectional diagram of a 10 gas turbine engine for an aircraft. The gas turbine engine 10 has a generally longitudinally extending centerline 12, extending from a forward direction 14 to a rear direction 16. The gas turbine engine 10 can include at least one counter-rotating portion. Therefore, the gas turbine engine 10 can? be defined as a counter-rotating gas turbine engine. The gas turbine engine 10 includes, in downstream series flow relationship, a fan section 18 including a front fan assembly 20 and a rear fan assembly 21, a counter-rotating compressor section 22 including at least one counter-rotating section, a combustion section 28 including a combustor 30, a counter-rotating turbine section 32 including at least one counter-rotating section, and an exhaust section 38.

[0018] In the illustrated gas turbine engine 10, the counter-rotating compressor section 22 can? comprise a counter-rotating low-pressure (LP) compressor 24 and a counter-rotating high-pressure (HP) compressor 26, while the counter-rotating turbine section 32 can include a counter-rotating HP turbine 34 and a counter-rotating LP turbine 36. Will this be understood? What aspects of disclosure may have applicability? with other turbine engines, including engines without counter-rotating sections, or turbine engines including a portion that is not ? counter-rotating. As a non-limiting example, aspects of disclosure may have applicability? with other gas turbine engines not including a counter-rotating LP turbine. For example, turbine engines having LP turbines in which circumferentially disposed static vanes are axially spaced apart from circumferentially disposed blades are also contemplated.

[0019] The fan assemblies 20 and 21 are positioned at one?end? front of the gas turbine engine 10 as shown. The terms ?front fan? and ?rear fan? are used herein to specify that one of the fan assemblies 20? coupled axially upstream of the rear fan assembly 21. ? It is also contemplated that the fan assemblies 20, 21 may be positioned at one end. rear of the gas turbine engine 10. The fan assemblies 20 and 21 each include a plurality of fan assemblies 20 and 21. of rows of fan blades 40 positioned within a fan housing 42. The fan blades 40 are joined to respective rotor discs 44 which are rotatably coupled through a respective front fan shaft 46 to the front fan assembly 20 and through a rear fan shaft 47 to the rear fan assembly 21.

[0020] The counter-rotating HP compressor 26, the combustor 30 and the counter-rotating HP turbine 34 form an engine center 48 of the gas turbine engine 10. The gas turbine engine center 48? surrounded by an external housing 50 which can? be coupled to the fan housing 42. The counter-rotating HP turbine 34 is coupled to the counter-rotating HP compressor 26 via a central rotor or shaft 52. During operation, the gas turbine engine center 48 generates combustion gases which are channeled downstream to the counter-rotating LP turbine 36 which extracts energy from the gases to power complexes of fan 20, 21 through their respective fan shafts 46, 47.

[0021] The counter-rotating LP turbine 36 includes an external rotor 54 positioned radially inward from the external housing 50. The external rotor 54 can have a generally frusto-conical shape and comprise a first series of aerodynamic surfaces 56, arranged circumferentially, which extend radially inwards towards the center line of the engine 12.

[0022] The counter-rotating LP turbine 36 further comprises an internal rotor 58 arranged substantially coaxially with respect to, and radially towards the interior of, the external rotor 54. The internal rotor 58 comprises a second series of aerodynamic surfaces 60 arranged circumferentially and spaced axially from the first series of aerodynamic surfaces 56. The internal rotor 58 can? furthermore be defined as a first rotor, while the external rotor 54 can be defined as a second rotor. The second set of aerodynamic surfaces 60 extend radially outwardly away from the engine centerline 12. The first and second sets of aerodynamic surfaces 56, 60 together define a plurality of aerodynamic surfaces 60. of turbine supports 62. In the example of Figure 1, five turbine supports 62 are shown and it will be understood? what can? any number of media can be used. Furthermore, although the first set of airfoil surfaces 56 is illustrated as facing the second set of airfoil surfaces 60, the first and second sets of airfoil surfaces 56, 60 may be arranged in any suitable manner, including the first set of airfoil surfaces 56 being positioned behind the second series of aerodynamic surfaces 60.

[0023] Although the gas turbine engine 10 is described in the context of comprising an outer rotor 54 and a rotating inner rotor 58, ? It is further contemplated that either the first set of airfoil surfaces 56 or the second set of airfoil surfaces 60 may be comprised in, or be part of, a fixed stator within the gas turbine engine 10. In one example, the first set of aerodynamic surfaces 56 pu? form a series of circumferentially arranged static fins forming part of an external stator within the gas turbine engine 10, while the second series of aerodynamic surfaces 60? coupled to the rotatable internal rotor 58.

In another example, the second set of airfoil surfaces 60 can? being in the form of static vanes coupled to an internal stator within the gas turbine engine 10, with the first set of aerodynamic surfaces 56 being in the form of blades coupled to an external rotor.

[0024] Complementary to the outer rotor 54 and the inner rotor 58, the stationary portions of the gas turbine engine 10, such as the outer housing 50, are also referred to individually or collectively as the stator 63. Accordingly, the stator 63 can? refer to the combination of non-rotating elements throughout the gas turbine engine 10.

[0025] During operation, the air flow exiting the fan section 18? divided such that a portion of the air flow is channeled along a main flow path 15 into the counter-rotating LP compressor 24, which then delivers pressurized air 65 to the counter-rotating HP compressor 26, which further pressurizes the air. The pressurized air 65 from the counter-rotating HP compressor 26? mixed with fuel in combustor 30 and ignited, thereby generating combustion gas 66 along the main flow path 15. Part of the work ? extracted from these combustion gases 66 by the counter-rotating HP turbine 34, which drives the counter-rotating HP compressor 26. The combustion gases 66 are disposed along the main flow path 15 into the counter-rotating LP turbine 36, which extracts additional work to drive the compressor counter-rotating LP 24, and the exhaust gas ? finally disposed of by the gas turbine engine 10 via the exhaust section 38. The drive of the counter-rotating LP turbine 36 can? activate the rotation of the front fan assembly 20 and the counter-rotating LP compressor 24.

[0026] A portion of the pressurized air 65 can be sucked in by the counter-rotating compressor section 22 as purge air 67. The purge air 67 can? be drawn in by pressurized air 65 and supplied to engine components requiring cooling. The temperature of the pressurized air 65 entering the combustor 30 ? increased significantly above the purge air temperature 67. The purge air 67 can? be used to reduce the temperature of the central components downstream of the combustor.

[0027] Part of the air delivered by the fan 20, such as the purge air 67, can bypass the center of gas turbine engine 48 and be used to cool portions, particularly hot portions, of the gas turbine engine 10, or to cool or power other portions of the gas turbine engine 10. In the context of an engine turbine, the hot portions of the gas turbine engine are normally downstream of the combustor 30, particularly the counter-rotating turbine section 32, with the counter-rotating HP turbine 34 being the ? the largest portion? hot since? ? directly downstream of the combustion section 28. Other sources of cooling fluid may be, but are not limited to, fluid disposed of by the counter-rotating LP compressor 24 or the counter-rotating HP compressor 26.

[0028] Figure 2 ? an axial cross-sectional view of the gas turbine engine 10 of Figure 1 as viewed from cut II of Figure 1. The gas turbine engine 10 can? comprise a rotor blade assembly 100 within a portion of the gas turbine engine 10. The rotor blade assembly 100 may comprise a rotor blade assembly 100. be provided inside a portion of the counter-rotating LP turbine 36. It will be observed? However, the rotating blade assembly 100 can be provided within any suitable portion of the gas turbine engine 10 as well as within any suitable portion of the counter-rotating compressor section 22 or the counter-rotating turbine section 32. Furthermore, although a single rotor blade assembly 100 is illustrated , will it be observed? what can? be any number of one or more? Rotating blade assemblies 100 provided within the gas turbine engine 10.

[0029] The rotating blade assembly 100 can? comprise a blade assembly 102, a disk 104, and a series of retainer assemblies 106. At least a portion of the disk 104 may be operatively coupled to a rotating component of the gas turbine engine 10. As a non-limiting example, the disk 104 can? be operatively coupled to a drive shaft 98 of the gas turbine engine 10. At least a portion of the blade assembly 102 may be operationally coupled to another rotating component. As a non-limiting example, the blade assembly can? be coupled to the internal rotor 58 or to the external rotor 54.

[0030] The blade assembly 102 can? comprising an internal platform 108, an external platform 110, located radially outward from the internal platform 108 with respect to the engine centerline 12, and a series of circumferentially spaced blades 112 extending therebetween. As illustrated, the array of circumferentially spaced blades 112 can understand the first series of aerodynamic surfaces 56. It will be observed? However, the series of circumferentially spaced blades 112 can include any suitable set of aerodynamic surfaces such as, but not limited to, the second set of aerodynamic surfaces 60. The set of circumferentially spaced blades 112 may be any blade or fin suitable inside the gas turbine engine 10 which is operatively coupled to the outer rotor 54, the inner rotor 58, or a static portion of the gas turbine engine 10 (e.g., the stator 63). The external platform 110 can? be operatively coupled to a rotary member of the gas turbine engine 10. As a non-limiting example, the external platform 110 can? be operatively coupled to the inner rotor 58 or outer rotor 54 of the gas turbine engine 10. A dovetail 114 may extend from a radially internal portion of the internal platform 108.

[0031] Disc 104 can? extend between external peripheries in the axial and radial directions. Disc 104 can? furthermore enclose or oppose at least a portion of the blade assembly 102. As a non-limiting example, the disk 104 can? enclose a radially internal portion of the blade assembly 102. As a non-limiting example, the disk 104 can? enclose a portion of the dovetail 114. The disk 104 can? also include internal suburbs. At least a portion of the internal suburbs can? oppose or come into contact with the assembly of blades 102 or the assembly of retaining elements 106. At least a portion of the internal peripherals can? define a seat 126 at least partially enclosing or opposing the dovetail 114. The seat 126 can? furthermore be defined by a first band 122 and a second band 124. As illustrated, the second band 124 can? be supplied axially forward or otherwise upstream, with respect to the combustion gases 66, of the first band 122. The disk 104 can? be sized and/or shaped such that the disc 104 can fit over, or otherwise enclose, at least a corresponding portion of the dovetail 114.

[0032] Disc 104 can? further include a projection 142 extending from a portion of the remainder of the disk 104. As illustrated, the projection 142 may extend from a portion of the first band 122. The projection 142 can? be operatively coupled to the drive shaft 98 of the gas turbine engine 10. The drive shaft 98 may be any suitable drive shaft 98 as described herein such as, but not limited to, the front fan shaft 46, the rear fan shaft 47, or the center shaft 52. Accordingly, in the event that the rotary blade assembly 100 is provided within the counter-rotating turbine section 32, the rotation of the rotating blade assembly 100 can? be used to rotate the drive shaft 98, which in turn can operating an upstream portion of the gas turbine engine 10 (e.g., a rotating component of the counter-rotating compressor section 22, a portion of the fan section 18, etc.). Although the projection 142 is illustrated, it will be noted that disk 104 can? be formed without the protrusion 142. Consequently, in some implementations the disk 104 is coupled only to the rotating blade assembly 100 and not to the drive shaft.

[0033] The assembly of retaining elements 106 can? extend through a portion of the disk 104 and operatively couple the disk 104 to the blade assembly 102. As illustrated, the retainer assembly 106 can extend axially through the disk 104 and abut the extremities? axially opposite of the disk 104. Consequently, the assembly of retaining elements 106 can? hold the disc 104 axially around a portion of the blade assembly 102. As a non-limiting example, the retainer assembly 106 can hold the disk 104 axially around the dovetail 114 of the blade assembly 102.

[0034] The blade assembly 102 can? be included within a series of circumferentially spaced blade assemblies 102. The series of blade assemblies 102 can extend around the entirety of the engine centerline 12 to form a ring of blade assemblies 102. At least a portion of the disk 104, however, may extend continuously across the entirety of the engine centerline 12. In other words, the disk 104 can form a 360 degree ring around the centerline of engine 12. Accordingly, the disk 104 can extend through one or more? blade assemblies 102 in the series of blade assemblies 102. Similarly, the projection 142 can? be formed as a continuous ring or band extending around the entirety of the engine centerline 12. Alternatively, the projection 142 may be formed as a continuous ring or band extending around the entirety of the engine centerline 12. be formed in distinct sections such that the projection 142 is included within a series of segmented projections 142 extending from respective portions of the disc 104. Accordingly, the disc 104 can be formed as a hub and rocket assembly when viewed in a plane normal to the engine centerline 12 and intersecting the disk 104. The array of retainer assemblies 106 may be formed as a hub and rocket assembly. comprise any suitable number of retainer assemblies 106 spaced circumferentially along the disk 104. As a non-limiting example, the array of retainer assemblies 106 may comprise any suitable number of retainer assemblies 106. be spaced evenly or otherwise circumferentially equal around the disk 104. Alternatively, the two or more? holding element assemblies 106 may be formed into groups in which the holding element assemblies 106 are more closer to each other than an adjacent group of holding element assemblies 106. In any case, the disk 104 can be held axially on each dovetail 114 of each blade assembly 102 of the set of blade assemblies 102 via the set of retainer assemblies 106.

[0035] As illustrated, the blade assembly 102, the disk 104, and the retainer assembly 106 are distinct components that are operatively coupled to, or otherwise opposed to, each other. It will be noted, however, that at least a portion of the blade assembly 102, disk 104, or retainer assembly 106 may be formed integrally with another portion of the rotating blade assembly 100. As a non-limiting example, the retaining element assembly 106 can be formed. be formed integrally with the second band 124 or with the first band 122 of the disk 104. As another non-limiting example, at least one of the first band 122 or the second band 124 can? be formed integrally with the blade assembly 102 such that the blade assembly 102 and the disk 104 form a monolithic structure.

[0036] Figure 3 ? is an exploded perspective view of the rotary blade assembly 100 of Figure 2. The rotary blade assembly 100 can? include the blade assembly 102, the disk 104 and the retainer assembly 106.

[0037] Each blade 112 of the series of circumferentially spaced blades 112 can understand to be defined by an external wall delimited by a front edge 116 and a rear edge 118, downstream or otherwise axially posterior to the front edge 116, a root 119 and a tip 120. The extension of the external wall between the front edge 116 and the rear edge 118 can? define a chord direction. The extension of the external wall between the root 119 and the tip 120 can define a width direction. The root 119 can? be coupled to, or otherwise integral with, the internal platform 108. The tip 120 can? be coupled to, or otherwise integral with, the external platform 110.

[0038] A first through hole 128 can? be defined by a portion of the disk 104. As a non-limiting example, the first band 122 and the second band 124 may each include a portion of the first through hole 128. A radially internal portion, or an end? distal of the dovetail 114, can? further comprise a portion of the first through hole 128. When assembled, the first through hole 128 can extend continuously through the first band 122, the dovetail 114 and the second band 124. The first through hole 128 can? extend from an upstream to a downstream portion of the rotating blade assembly 100 in the axial direction.

[0039] Each holding element assembly 106 of the holding element assembly series 106 can? comprise a tubular element, a pin 132 and a fastening element 134. The tubular element can? include any suitable tubular shape when viewed in a plane normal to the engine centerline 12 and intersecting the tubular member such as, but not limited to, a square tube, a cylindrical tube or any other suitable tube. At least a portion of the holding element assembly 106 can abut, be in contact with, or be coupled to the seat 126 of the disc 104.

[0040] As a non-limiting example, the tubular element can? be any tubular element suitable as a bushing 130. The bushing 130 can? comprise a series of fingers 131 extending along a portion of the bushing 130. Each finger 131 of the series of fingers 131 can? be separated from an adjacent finger 131 by a void or absence of material. The bushing 130 can? comprise a hollow interior defining a second through hole 136. Consequently, the tubular element can also be defined as a hollow tubular element. The pin 132 can? be aligned with, and received at least partially inside, the first through hole 128 and the second through hole 136. The pin 132 can? terminate at one?extremity? distal 138. When assembled, the end? distal 138 can? extend beyond the first through hole 128. The bushing 130 can? be aligned with the first through hole 128. The bushing 130 can? also include a first extremity? defining a shoulder 140 which, when assembled, can? abut on at least a portion of the disk 104. As a non-limiting example, the shoulder 140 can? rest on the second band 124. The fastening element 134 can be secured at the end? distal 138 of the pin 132 and rest on a portion of the disk 104. The tubular element can be defined generally as any suitable element that ? expandable through mechanical characteristics (e.g. fingers 131) when an external force is applied to an internal part of the tubular element. ? it is also contemplated that the tubular element can be expandable through properties? of material (for example thermal properties, elasticity, etc.). As a non-limiting example, the tubular element can be a rubber tube that expands when external force ? applied by pin 132.

[0041] Figure 4 ? a cross-sectional view of the rotary blade assembly 100 viewed from cut IV-IV of Figure 3.

[0042] The first band 122 can? include a first rib 144. The second band 124 can? comprise a second rib 146, opposite the first rib 144. As illustrated, the first rib 144 and the second rib 146 may be provided on axially opposite sides of the dovetail 114. The first rib 144 and the second rib 146 may both interface, or otherwise contact, a corresponding portion of the dovetail 114. The first rib 144 and the second rib 146 may be used to radially retain the disk 104 on the blade assembly 102.

[0043] The second band 124 can? comprise a portion that overlaps with a corresponding portion of the first band 122. This overlapping portion can? define an overlap junction 145 formed between the first band 122 and the second band 124. The overlap junction 145 can? define a coupling or an interface between the first band 122 and the second band 124. ? the possibility contemplated? that the overlap junction 145 can also be used to align the first band 122 with respect to the second band 124.

[0044] The second band 124 can? comprise a portion that overlaps with a corresponding portion of the first band 122. This overlapping portion can? define the overlap junction 145 formed between the first band 122 and the second band 124. The overlap junction 145 can? define a coupling or an interface between the first band 122 and the second band 124. ? the possibility contemplated? that the overlap junction 145 can also be used to align the first band 122 with respect to the second band 124.

[0045] As illustrated, the bushing 130 of the holding element assembly can extend through a portion of the first through hole 128 such that the second through hole 136 is aligned with the first through hole 128. ? contemplated that the bushing 130 could end along one end? distal 148 of the bushing 130. The extremity? distal 148 of the bushing 130 pu? be provided inside a portion of the first through hole 128. As a non-limiting example, the end? distal 148 can? be spaced from the first band 122 in such a way that the bush 130 does not physically come into contact with the first band 122.

[0046] ? It is further contemplated that the first through hole 128 and the second through hole 136 may each comprise a non-constant cross-sectional area when viewed in a plane normal to the engine centerline 12 and intersecting the disk 104. As a non-limiting example, the first through hole 128 can? comprise an area with a reduced cross-sectional area within a portion of the first band 122 when compared to the remainder of the first through hole 128. The portion of the first through hole 128 with the reduced cross-sectional area may ? come into direct contact with at least a portion of the pin 132. As a non-limiting example, the bush 130 can? comprise a portion with decreasing cross-sectional area. In other words, the bushing 130 can? include a portion in which the cross-sectional area decreases linearly or nonlinearly. As a non-limiting example, at least a portion of the cross-sectional area of the bushing 130 can decrease from an upstream or axially anterior portion to a downstream or axially posterior portion. Similarly, the pin 132 can comprise a decreasing cross-sectional area when viewed in a plane normal to the engine centerline 12 and intersecting the pin 132. Accordingly, the pin 132 can be defined as a conical pin and the bushing can? be defined as a conical bushing. The decreasing cross-sectional area of the pin 132 can correspond to the decreasing cross-sectional area of the bushing 130 such that the pin 132 can interface with the bushing 130 along the sections of the bushing 130 and the pin 132 defined by the decreasing cross-sectional areas. The interface between the bushing 130 and the pin 132 can be used to retain the pin 132 inside the bush 130.

[0047] The shoulder 140 of the bush 130 can? interface with the disk 104. As a non-limiting example, the shoulder 140 of the bushing 130 can? interface with the second band 124. The second band 124 can? include a cutout 150 sized to receive the shoulder 140 of the bushing 130. Consequently, a front portion of the bushing 130 can be flush with a front portion of the second band 124.

[0048] ? It is contemplated that at least a portion of the fastening element 134 can abut a portion of the first band 122. The fastening element 134 can be any suitable fastener 134 such as, but not limited to, a nut, a hydraulic fastener, a magnetic fastener, a weld, an adhesive or an electrical connection (e.g. an electronically activated connection). As a non-limiting example, the fastener 134 can be defined by a die. Consequently, the extremity distal 138 can? further comprising a threaded portion corresponding to a threaded portion of the nut. Consequently, the die can? be fixed or otherwise threaded, on the threaded portion of the pin 132.

[0049] During assembly, the disk 104 can be inserted over a corresponding portion of the dovetail 114 such that the first through hole 128 is formed continuously through the first band 122, the second band 124 and the dovetail 114. The first rib 144 and the second rib 146 can each interface with a corresponding portion of the dovetail 114. The overlap junction 145 can? be sized and positioned to ensure that the first rib 144 and second rib 146 are positioned within the correct position when assembled. Furthermore, the overlap junction 145 can be sized to ensure that the first through hole 128 is formed continuously through the disk 104 and the dovetail 114 once the disk 104 is ? positioned above the dovetail 114. The bushing 130 can? be subsequently aligned with the first through hole 128 and inserted inside it. The bushing 130 can? be inserted in such a way that the shoulder 140 comes into contact with or is received inside the cutout 150. The pin 132 can? then be inserted into the second through hole 136 defined by the bushing 130. As discussed herein, the end? distal 138 of the pin 132 pu? extend beyond a termination of the first through hole 128. The fastening element 134 can then be placed, applied, fixed or otherwise coupled to the extremity? distal 138 of the pin 132 and abut on a portion of the disk 104 (for example the first band 122). The fastener 134 can apply an axial tightening or closing force on the pin 132 to pull the pin 132 towards the fastener 134, which is constrained axially by the disk 104. When the pin 132 is ? pulled towards the fastener 134, the pin 132 is first moved axially until? Not ? in contact with the bush 130, where the continued axial movement of the pin 132 subsequently causes that the shoulder 140 abuts on the cutout 150 formed inside the disc 104. Through the shoulder 140 abutting the cutout 150 and the fastening element 134 abutting the disc 104, opposite closing forces are exerted on the ends? opposite axial forces of the disk 104, which in turn axially hold the disk 104 above the dovetail 114. Any additional axial movement of the pin 132 causes s? that the fingers 131 of the bushing 130 expand radially and apply a loop-like expansive force between the disk 104 and the dovetail 114. Consequently, the bushing 130 can further be defined as an expandable bushing or an expandable tubular member, respectively, understanding that the expansion of the bushing 130 can be generated via any suitable method such as a mechanical component, or a property? of material of the bushing 130. Similarly, the pin 132 can? be defined generally as a component configured to activate and expand the bushing 130 as described herein. The ring-shaped expansive force, in turn, pushes the dovetail 114 against the first rib 144 and the second rib 146. Therefore, with this type of connection, the assembly of retaining elements 106 axially constrains the disk 104 on the dovetail swallow 114 as well as? radially constrains the disk 104 on the dovetail 114.

[0050] During operation of the gas turbine engine 10, a flow of working air can flow over a portion of the rotary blade assembly 100. As a non-limiting example, the work air flow may flow over the series of circumferentially spaced blades 112 of the rotating blade assembly 100. As a non-limiting example, the flow of work air can be any suitable air flow within the gas turbine engine such as, without limitation, pressurized air 65 or combustion gases 66. Where the rotor blade assembly 100 is provided within the turbine section counter-rotating blade 32, the rotating blade assembly 100 can? extract work from the work air stream as it flows over the rotor blade assembly 100. In the event that the rotor blade assembly 100 is provided within the counter-rotating compressor section 22, the rotor blade assembly 100 can pressurize or otherwise compress the working air flow.

[0051] As discussed herein, the rotary blade assembly 100 can comprising a series of circumferentially spaced blade assemblies 102 that are distinct from one another. ? It is contemplated that the disk 104 may be used to couple each blade assembly 102 of the series of blade assemblies 102 to form a continuous ring of blade assemblies 102. Accordingly, the rotating blade assembly 100 may be formed as a rigid rotating blade assembly 100 by connecting the blade assemblies 102 together through the use of the disk 104. This, in turn, can be formed. ensure that there is no radial or otherwise minimal clearance between the external rotor 54 and the outer ring 110, and a portion of the disc 104 (e.g. the projection 142) and the drive shaft 98. Reducing or eliminating the clearances can ensure that the rotation of the blade array 102 within the rotary blade assembly 100 is concentric with the rotation of the external platform 110 or external rotor 54. Similarly, reducing or eliminating clearances can ensure that the disk 104 is concentric with the outer band 110. With the concentric rotation and the assembly of the rotating blade assembly 100, the quantity? total losses? reduced (e.g. friction losses through adjacent pieces abutting each other. This ultimately ensures that the overall efficiency of the gas turbine engine 10 is increased when compared to the gas turbine engine 10 if it did not include the overall of rotating blades 100 with the disk 104.

[0052] During normal operation of the gas turbine engine 10, on the rotating blade assembly 100 can? an operational force be exerted. The operational force can be defined as any suitable force exerted on the rotating blade assembly 100 during normal operation of the gas turbine engine 10 (e.g., a rotational force or a thermal load). As a non-limiting example, on the rotating blade assembly 100 can? an operating force of 30 klb can be exerted. The operational force can be exerted on the disk 104 and define a force radially inwards with respect to the center line of the engine 12. ? It is contemplated that the disc 104 may be formed to resist these operating forces of the gas turbine engine 10. As a non-limiting example, the interface of the disc 104 with the dovetail 114 (e.g., the first rib 144 and the second rib 146) can? be sized or formed to withstand these operational forces. During the shutdown of the gas turbine engine 10, however, on the rotating blade assembly 100? a stopping force is exerted, opposite to the operating force. In other words, during the shutdown of the gas turbine engine 10, on the rotating blade assembly 100 can a force be exerted radially outwards. The stopping force can? be smaller than the operational forces. As a non-limiting example, the stopping force can? be 1/20<or>of the operational force. As a non-limiting example, if the operational force is? of 30 klb, the stopping force can? be 1.5 klb. Unlike the operational force, however, the stopping force is ? transferred through a portion of the holding element assembly 106 instead of? only to the disk 104. Consequently, the set of retaining elements 106 is sized and formed to withstand the crash force. Since? the stopping force? much more? small of the operational force, the complex of holding elements 106 can? be formed with a more material? weaker than the disk 104, which ultimately reduces the material costs associated with the rotary blade assembly 100.

[0053] Figure 5 ? a cross-sectional view of an exemplary rotor blade assembly 200 for use within the gas turbine engine of Figure 1. The exemplary rotor blade assembly 200 is similar to rotary vane assembly 100; therefore, similar parts will be identified with equal numbers in the series 200 , understanding that the description of the similar parts of the rotary blade assembly 100 applies to the exemplary rotary blade assembly 200 , unless otherwise indicated.

[0054] The rotating blade assembly 200 can? comprise a blade assembly 202, a disk 204, and a retainer assembly 206 similar to the rotating blade assembly 100. The blade assembly 202, similar to the blade assembly 102, may comprise a blade 212 extending from a root 219 to a tip (not shown) and from a leading edge 216 to a trailing edge 218. The root 219 may be coupled to an internal platform 208 of the blade assembly 202, while the tip can? be mated to an external platform (not shown) of the blade assembly 202. A dovetail 214 can? depend on the internal platform 208. The disk 204, similar to the disk 104, can? include a first band 222 and a second band 224 which together form a seat 226. The first band 222 can? optionally be operatively coupled to a drive shaft through a protrusion 242. The retainer assembly 206, similar to the retainer assembly 106, may be comprise a bushing 230, a pin 232 and a fastener 234. The bushing 230 can? include a shoulder 240 and a series of fingers 231, which interfaces with a cutout 250 formed inside a portion of the second band 224. The pin 232 can be defined by an?extremity? distal 238 and the fixing element 234 can? be secured at the end? distal 238. At least a portion of the disk 204 and the dovetail 214 can? form a first continuous through hole 228. An internal part of the bushing 230 can? define a second through hole 236 aligned with the first through hole 228. The pin 232 can? be provided at least partially inside the second through hole 236 and the first through hole 228. An overlap junction 245 can? be formed between the first band 222 and the second band 224 and define an interface or coupling between the first band 222 and the second band 224.

[0055] Disc 204 ? similar to disc 104 since? includes the first band 222 and the second band 224. The first band 222, however, forms a plate abutting a portion of the dovetail 214 and the second band 224. The second band 224 can however, be formed to extend only through a radially inner portion of the rotary vane assembly 200. In other words, the second band 224 does not extend in front of, or otherwise abut, a portion of the dovetail 214 along a portion of the assembly of rotating blades 200 including the assembly of retaining elements 206. Furthermore, only the portion of the first through hole 228 defined by the second band 224 is the portion of the second band 224 opposite the dovetail 214. In other words, the second band 224 does not define a complete portion of the first through hole 228 on its own as the second band 124 does (for example the hole formed inside a axially anterior portion of the second band 124). The bushing 230 can? extend through a portion of the first through hole 228 and terminate within the first through hole 228 at a termination 248.

[0056] The cutout 250 ? similar to the 150 cutout, however, the 250 cutout is also formed at least partially inside the dovetail 214. Consequently, the shoulder 240 of the bushing 230 can? abut at least a portion of the disk 204 and the dovetail 214 or on the blade assembly 202. As illustrated, the first through hole 228 has a shorter axial length than the first through hole 128. This? ? due to disk configuration 204.

[0057] The bush 230 ? similar to the bushing 130, except that the bushing 230 can? have a shorter axial length when compared to the bushing 130. There? ? due to the shorter axial length of the first through hole 228. This, in turn, reduces the material required for the fabrication of the retainer assembly 206. Furthermore, the bushing 230 can include a wall 251, which ends at the extremities? radially distal to define the shoulder 240.

[0058] During the operation of the gas turbine engine 10, at least a portion of the working air flow can flow towards the disk 204, thus defining a leakage fluid. ? It is contemplated that minimizing the leakage fluid within the gas turbine engine 10 can minimize the amount of of working air flow flowing over the blades 212, which in turn maximizes the amount? of working air extracted from the working air flow. With the wall 251 and the shoulder 240, the bushing 230 can? form a fluid-tight seal between the disk 204 and the blade assembly 202. This, in turn, ensures that an escaping fluid does not enter either the first through hole 228 or the second through hole 236. This reduces the quantity? total escaping fluid, which in turn maximizes the overall efficiency of the rotary blade assembly 200. ? contemplated that the remaining portion of the disk 204 may be used to limit the leakage fluid. As a non-limiting example, the first band 222 can? be used to reduce or otherwise eliminate leakage fluid, which can flow from an upstream portion of the rotor blade assembly 200 to a downstream portion of the rotor blade assembly 200.

[0059] Figure 6 ? a radial view of the rotor blade assembly 200 of FIG. 5 viewed from a plane normal to the engine centerline 12 and intersecting the rotor blade assembly 200. The rotor blade assembly 200 may include the blade assembly 202, which is comprised within a series of blade assemblies 202 spaced circumferentially with respect to each other.

[0060] Disc 204 can? be mated to the dovetail 214 through a dovetail connection defined by a tail 254 and a socket 256. The disk 204 can be mated to the dovetail 214. include the tail 254, which extends radially, with respect to the center line of engine 12, from a remaining part of the disk 204. As a non-limiting example, the second band 224 can? include the tail 254, which extends radially from the remaining part of the second band 224.

[0061] The socket 256 can? be formed between circumferentially adjacent portions of adjacent blade assemblies 202. As illustrated, socket 256 can? be formed by circumferentially adjacent cutouts formed within a portion of adjacent dovetails 214. Can socket 256? be sized and shaped to receive the tail 254 of the disk 204.

[0062] During the assembly of the rotating blade assembly 200, the tail 254 of the disk 204 can be inserted through or into the cavity? 256. There? can? be carried out by sliding the tail 254, and therefore the disk 204, into the socket 256. At least a portion of the tail 254 can? interface with the socket 256. The interface between the queue 254 and the socket 256 can? radially retain the disk 204 on the blade assembly 202. This radial retention through the tail 254 and the socket 256 is similar to the radial retention between the first rib 144 and the second rib 146, and the dovetail 114 of the rotary blade assembly 100. Furthermore, since? the bushing 230 comes into direct contact with the dovetail 214, at least a portion of the closing force that holds the disk 204 axially on the blade assembly 202 can? be applied directly to the blade assembly 202.

[0063] Figure 7 ? a cross-sectional view of an exemplary rotor blade assembly 300 of the gas turbine engine 10 of Figure 1. The exemplary rotor blade assembly 300 is similar to rotary blade assembly 100, 200; therefore, similar parts will be identified with equal numbers in the series 300, understanding that the description of the similar parts of the rotary blade assembly 100, 200 applies to the exemplary rotary blade assembly 300, unless otherwise indicated.

[0064] The rotating blade assembly 300 can? comprise a blade assembly 302, a disk 304, and a retainer assembly 306 similar to the rotating blade assembly 100, 200. The blade assembly 302, similar to the blade assembly 102, 202 may comprise a blade 312 extending from a root 319 to a tip (not shown) and from a leading edge 316 to a trailing edge 318. The root 319 may be coupled to an internal platform 308 of the blade assembly 302, while the tip can? be mated to an external platform (not shown) of the blade assembly 302. A dovetail 314 can? depend on the internal platform 308. The disk 304, similar to the disk 104, 204, can? define a seat 326. The disk 304 can? optionally be operatively coupled to a drive shaft through a protrusion 342. The holding element assembly 306, similar to the holding element assembly 106, 206, can be comprise a bushing 330, with a set of fingers 331, a pin 332, and a fastener 334. The bushing 330 can? be formed in a similar way to the bushing 230 since? includes a wall 351 which terminates at the extremities radially distal to define a shoulder 340. Shoulder 340 can? interface with a cutout 350 formed at least partially within a portion of the dovetail 314 and a portion of the disk 304. The bushing 330 can furthermore be defined by a shorter axial length, similar to the bushing 230, when compared to the bushing 130. The pin 332 can be defined by an?extremity? distal 338 and the fixing element 334 can? be secured at the end? distal 338. At least a portion of the disk 304 and the dovetail 314 can? form a first continuous through hole 328. An internal part of the bushing 330 can? define a second through hole 336 aligned with the first through hole 328. The pin 332 can? be provided at least partially inside the second through hole 336 and the first through hole 328. The bush 330 can? extend through a portion of the first through hole 328 and terminate within the first through hole 328 at a termination 348.

[0065] Disc 304 ? similar to disc 204 in that it is formed to extend only through a radially internal portion of the rotary vane assembly 200. In other words, the disc 204 does not contact or interface with an axially forward portion of the dovetail 314. The difference between the disc 304 and the disc 204 ? that the disk 204 includes the first band 222 and the second band 224. The disk 204 can? however, be defined as a solid disk 304 in which the first band 122, 222? formed in solidarity with the second band 124, 224. In other words, disk 304 is formed as a monolithic structure that is held axially on the blade assembly 302 via the holding element assembly 306. The disk 304 can? further be held radially through the use of any suitable radial holding assembly as described herein (e.g. tail 254 and socket 256, or first rib 144 and second rib 146).

[0066] Figure 8 ? a cross-sectional view of an exemplary rotor blade assembly 400 of the gas turbine engine 10 of Figure 1. The exemplary rotor blade assembly 400 is similar to the rotating blade assembly 100, 200, 300; therefore, similar parts will be identified with equal numbers in the series 400, understanding that the description of the similar parts of the rotary blade assembly 100, 200, 300 applies to the exemplary rotary blade assembly 400, unless otherwise indicated.

[0067] The rotating blade assembly 400 can? comprise a blade assembly 402, a disk 404, and a retainer assembly 406 similar to the rotating blade assembly 100, 200, 300. The blade assembly 402, similar to the blade assembly 102, 202, 302 can comprise a blade 412 extending from a root 419 to a tip (not shown) and from a leading edge 416 to a trailing edge 418. The root 419 may be coupled to an internal platform 408 of the blade assembly 402, while the tip can? be mated to an external platform (not shown) of the blade assembly 402. A dovetail 414 can? depend on the internal platform 408. The disk 404, similar to the disk 104, 204, 304, can? define a seat 426. The disk 404 can? optionally be operatively coupled to a drive shaft through a protrusion 442. The disc 404 can? be formed in a similar way to disk 304 in that it? a supportive 404 disc. The holding element assembly 406, similar to the holding element assembly 106, 206, 306, can comprise a bushing 430, a pin 432 and a fastener 434. The bushing 430 can? be formed in a similar way to the bushing 130, 230, 330 as it includes a shoulder 440 and a series of fingers 431. The pin 432 can? be defined by an?extremity? distal 438 and the fixing element 434 can? be secured at the end? distal 438. At least a portion of the disk 404 and the dovetail 414 can? form a first continuous through hole 428. An internal part of the 430 bushing can? define a second through hole 436 aligned with the first through hole 428. The pin 432 can? be provided at least partially inside the second through hole 436 and the first through hole 428. The bush 430 can? extend through a portion of the first through hole 428 and terminate within the first through hole 428 at a termination 448.

[0068] The assembly of retaining elements 406 can? further comprise a retainer plate 458. The retainer plate 458 may abut on a portion of the dovetail 414 and the disc 404. The plate of retaining elements 458, similar to the disc 104, 204, 304 and the dovetail 214, 314, can? include a cutout 450 configured to receive the shoulder 440 of the bushing 430. The retainer plate 458 may further comprise a through hole 460 that extends axially through a portion of the retainer plate 458. The retainer plate 458 may be aligned with the first through hole 428 such that the through hole 460 defines a portion of the first through hole 428.

[0069] During the assembly of the rotating blade assembly 400, at least one closing force, or the axial holding force generated by the holding element assembly 406, can be applied to the retainer plate 458. Accordingly, the retainer plate 458 can be applied to the retainer plate 458. be used to axially retain the disk 404 on the blade assembly 402.

[0070] Figure 9 ? a schematic axial view of an exemplary rotor blade assembly 500 of the gas turbine engine 10 of Figure 1. The exemplary rotor blade assembly 500 is similar to the rotating blade assembly 100, 200, 300, 400; therefore, similar parts will be identified with equal numbers in the series 500, understanding that the description of the similar parts of the rotary blade assembly 100, 200, 300, 400 applies to the exemplary rotary blade assembly 500, unless otherwise indicated.

[0071] The rotating blade assembly 500 can? comprise a set of blade assemblies 502, a disk 504, and a set of retainer assemblies 506 similar to the rotary blade assemblies 100, 200, 300, 400. Each blade assembly 502 of the set of blade assemblies 502 may be included. be similar to the blade assembly 102, 202, 302, 402. Each blade assembly 502 can? comprising a blade 512 or a series of blades 512, with each blade 512 extending from a root 519 to a tip (not shown) and from a leading edge 516 to a trailing edge (not shown). The root 519 can? be coupled to an internal platform 508 of the blade assembly 502, while the tip can? be coupled to an external platform (not shown) of the blade assembly 502. Each blade assembly 502 can? include a 514 dovetail that can? depend on the corresponding internal platform 508 of the blade assembly 502. The disk 504, as illustrated, is a schematic illustration of disk 504. It will be observed? however, that the 504 disk can? be any suitable disk 104, 204, 304, 404 as disclosed herein. The set of retainer assemblies 506, as illustrated, is a schematic illustration of the retainer assembly 506. Please appreciate? However, each holding element assembly 506 of the holding element array 506 can? comprising any suitable retainer assembly 106, 206, 306, 406 as described herein.

[0072] As illustrated, the series of blade assemblies 502 can? each including a corresponding dovetail 514. Each 514 dovetail can? be formed so as to be complementary with an adjacent 514 dovetail. A first contact region 562 can? be formed between corresponding portions of adjacent dovetails 514. The first contact region 562 can? indicate a region where adjacent dovetails 514 physically contact each other or are otherwise coupled together. Furthermore, each 514 dovetail can? be touched by a portion of the disc 504. As a non-limiting example, each dovetail 514 can? be touched by the disk 504 along a second contact region 564. As a non-limiting example, the second contact region 564 can be the contact or interface between the first rib 144 or the second rib 146 of the rotating blade assembly 100. As a non-limiting example, the second contact region 564 can be be the contact or interface between the tail 254 and the socket 256 of the rotating blade assembly 200.

[0073] ? contemplated that the series of holding element assemblies 506 may be arranged in groups. As a non-limiting example, the series of holding element assemblies 506 can comprise a group of two holding element assemblies 506. It will be observed? However, each group of holding element assemblies 506 can understand any number of one or more? holding element assemblies 506. As illustrated, each group of holding element assemblies 506 of the set of holding element assemblies 506 can? extend through a portion of the disk 504 corresponding to each other blade assembly 502. In other words, each other blade assembly 502 can be physically coupled to a holding assembly 506 of the holding assembly series 506. This configuration may be physically coupled. reduce the total number of retainer assemblies 506 required in order to effectively couple the disk 504 to the array of blade assemblies 502. Alternatively, the retainer assemblies 506 may be distributed across any number of blade assemblies 502. As a non-limiting example, the series of holding element assemblies 506 can be provided along a portion of the disk 504 corresponding to every third, fourth, fifth or nth blade assembly 502. Alternatively, the series of retainer assemblies 506 may be provided along a portion of the disk 504. be provided along the disk 504 corresponding to each blade assembly 502 in the series of blade assemblies 502.

[0074] The benefits of the present disclosure include a rotating blade assembly that can be used with a turbine engine (e.g. a counter-rotating turbine engine), which has improved overall efficiency when compared to a conventional turbine engine (e.g. a non-counter-rotating turbine engine). For example, conventional turbine engines may include a set of rotating blades provided downstream of the set of stationary blades, which together form a non-counter-rotating turbine engine mount. During operation of the non-counter-rotating turbine engine, a flow of working air, similar to the flow of working air as described herein, may flow over the set of stationary blades and then to the set of adjacent rotating blades. The series of stationary fins can? be used to direct the flow of working air to be incident with that leading edge of the rotating blade array, thereby limiting the air resistance losses associated with a non-incident working air flow. In a non-counter-rotating turbine engine, however, work ? extracted only through the rotation of the rotating blade array (e.g. the stationary blade array does not extract work from the working air stream). The present disclosure, however, concerns a rotating blade assembly for a counter-rotating turbine engine in which a support is consisting of two adjacent rotating blade complexes. ? It is contemplated that adjacent rotor blade assemblies may rotate in opposite (e.g. opposite) circumferential directions to each other, however the rotor blade assemblies may be positioned such that the flow of working air exiting the rotor blade assembly rotating blades upstream may be impinging on their leading edges. As a result, air resistance losses are still avoided or otherwise limited, however work can be done. be extracted from both sets of rotating blade assemblies. There? finally, it means that the total work emission from the counter-rotating turbine engine can be greater when compared to a non-counter-rotating turbine engine of similar size (e.g. similar or equal amount of total mounts).

[0075] Additional benefits of the present disclosure include a rotor blade assembly with reduced losses when compared to a rotor blade assembly used within a counter-rotating turbine engine without the disc as described herein. For example, a rotating blade assembly without the disc as described in this former? a non-rigid circumferential ring of blade assemblies within the working air flow. This, in turn, means that the blade assemblies have greater tolerances regarding how far they can move during the desired circumferential displacement (e.g., rotation) of the rotating blade assembly. The larger tolerances, in turn, make s? that the internal portions of the rotating blade assembly and the outer band are non-concentric, which ultimately generates losses. The rotary blade assembly, as described herein, however, includes a circumferential disk which interconnects each of the blade assemblies within the rotary blade assembly. In other words, the disk can? be used to form a rigid structure between adjacent blade assemblies. This, in turn, reduces or eliminates clearances between adjacent components, which ensures concentricity. between the disc and the external platform as described herein. The concentric rotation and assembly of the rotary blade assembly relative to the outer shroud, in turn, minimizes the losses that are generated, which ultimately increases the efficiency of the rotary turbine engine when compared to a conventional rotary turbine engine without the rotating blade assembly as described herein.

[0076] Additional benefits of the present disclosure include a rotating blade assembly within the counter-rotating turbine engine having capacity increased friction damping when compared to a conventional rotary blade assembly used within a counter-rotating turbine engine. For example, conventional rotary blade assemblies may use a monolithic structure interconnecting adjacent blade assemblies. In other words, conventional rotary blade assemblies may comprise a single inner shroud, formed as a single unitary piece that is integral with the remaining part of the rotating blade assembly. The monolithic structure, however, has the capacity friction damping low since? aren't you? a contact surface between adjacent components. As a result, the conventional rotor blade assembly will vibrate, which in turn increases the total losses associated with the operation of the conventional counter-rotating turbine engine. The counter-rotating turbine engine, as described herein, however, comprises a non-monolithic assembly of rotating blades. As a non-limiting example, the counter-rotating turbine engine, as described herein, can understand a non-monolithic disk that ? coupled to and contacts the remainder of the rotor blade assembly along various contact regions generated by the interface between the remainder of the rotor blade assembly and the disk (e.g., the fastener assembly, ribs, and tails/sockets). These contact regions can create areas of frictional contact between two adjacent portions in the rotating blade assembly. This, in turn, increases the capabilities friction damping of the rotor blade assembly when compared to the conventional rotor blade assembly. There? reduces vibration losses associated with the operation of the counter-rotating turbine engine when compared to the conventional counter-rotating turbine engine, which ultimately increases the overall efficiency of the counter-rotating turbine engine when compared to the conventional counter-rotating turbine engine.

[0077] To the extent not already? described, the different characteristics and structures of the various aspects can be used in combination with each other as desired. The fact that a feature may not be illustrated in all aspects does not mean that it may not exist, but does it? ? done for brevity? of description. Therefore, the various characteristics of different aspects can be mixed and matched as desired to form new aspects, whether the new aspects are specifically described or not. Combinations or permutations of features described herein are covered by this disclosure.

[0078] This written description uses examples to describe aspects of the disclosure described herein, including how? best, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of disclosure aspects? defined by the claims and can? understand other examples that may come to mind for those skilled in the art. Such other examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

[0079] Further aspects of disclosure are provided by the topic of the following main points.

[0080] A rotor blade assembly for a turbine engine having a drive shaft, the rotor blade assembly including a disk operatively coupled to the drive shaft and including a seat having at least a portion of a first through hole, at least one assembly of blades having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform and a retainer assembly securing the disk to the blade assembly, the retainer assembly retention including a hollow tubular element defining a second through hole, which is aligned with the first through hole, and a pin extending through at least a portion of the second through hole and the first through hole and abutting at least a portion of the hollow tubular member.

[0081] The rotating blade assembly of any previous principal point, in which the hollow tubular element is an expandable bushing.

[0082] The rotating blade assembly of any previous principal point, where the pin terminates at an end? distal and the rotary blade also includes a fastener secured to the end? distal and abutting the disc.

[0083] The rotating blade assembly of any previous major point, where the fastener is? one of a nut, a hydraulic fastener, a magnetic fastener, a weld, an adhesive, or an electrical connection.

[0084] The rotating blade assembly of any previous major point, where the fastener is? the nut and stud includes a threaded section, where the nut is threaded onto the threaded section of the pin.

[0085] The rotating vane assembly of any previous principal point, where the dovetail defines at least one other portion of the first through hole.

[0086] The rotating blade assembly of any preceding principal point, wherein the disk further includes a first band and a second band which together define the seat.

[0087] The rotating blade assembly of any preceding principal point, wherein the first band and the second band are coupled at an overlap junction.

[0088] The rotary vane assembly of any preceding principal point, wherein the first band comprises a first rib and the second band comprises a second rib, both the first rib and the second rib abutting a corresponding portion of the dovetail.

[0089] The rotary vane assembly of any foregoing principal, wherein the retainer assembly further includes a retainer plate including a through hole aligned with the first through hole and abutting a portion of the disk and tail of swallow.

[0090] The rotary blade assembly of any preceding principal point, wherein the hollow tubular member further includes a first end? comprising a shoulder, with the shoulder abutting a corresponding portion of the retainer plate.

[0091] The rotating vane assembly of any preceding principal, wherein the hollow tubular member further includes a shoulder abutting at least one of the disc or dovetail.

[0092] The rotating blade assembly of any preceding principal point, wherein the disk includes a dovetail link extending from a remainder of the disk in the width direction and extending through a corresponding portion of the dovetail.

[0093] The rotating blade assembly of any previous major point, where the disk extends continuously around the drive shaft 360 degrees.

[0094] A gas turbine engine, including an engine center defining an engine centerline and including a drive shaft and a first rotor, and a rotating blade assembly, including a disk operatively coupled to the drive shaft and comprising a seat having at least a portion of a first through hole, at least one blade assembly having an upper platform operatively coupled to the first rotor, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the platform lower, and at least one assembly of retaining elements securing the disk to the at least one assembly of blades, the at least one assembly of retaining elements including a hollow tubular element defining a second through hole, which is aligned with the first through hole, and a pin extending through at least a portion of the second through hole and the first through hole and abutting at least a portion of the hollow tubular member.

[0095] The gas turbine engine of any preceding principal point, wherein the at least one blade assembly is comprised within a series of blade assemblies spaced circumferentially apart from each other and extending circumferentially around the entirety of the engine centerline, and in which the disk is a 360-degree ring that extends circumferentially around each dovetail in the series of blade assemblies.

[0096] The gas turbine engine of any preceding principal point, wherein the at least one retainer assembly is comprised with a set of holding member assemblies and wherein the set of holding member assemblies is provided along the disk at circumferential locations corresponding to each other blade assembly and wherein at least two adjacent blade assemblies define a socket extending at least partially through each dovetail and wherein the disk further includes a tail extending from the remainder of the disc and received at least partially inside the socket.

The gas turbine engine of any foregoing principal, further including an outer rotor spaced radially outward from the first rotor relative to the engine centerline.

[0098] The gas turbine engine of any foregoing principal, wherein the disk further includes a first band comprising a first rib abutting the dovetail, the first band defining a first portion of the seat, and a second band comprising a second rib opposing the dovetail, the second band defining a second portion of the seat in which the first rib and the second rib radially retain the disc on the dovetail.

[0099] The gas turbine engine of any previous main point, in which the hollow tubular element is an expandable bushing, the expandable bushing further including a shoulder abutting at least one of the disc or dovetail, and in which the pin terminates at an end distal, the at least one retainer assembly further includes a retainer plate comprising a through hole aligned with the first through hole and abutting a portion of the disc and dovetail, and at least one fastener secured to the ?extremity? distal and abutting the disc.

Part # Part

10 gas turbine engine

12 Engine centerline

14 Front direction

15 Main flow path

16 Rear direction

18 Fan section

20 Front fan assembly

21 Rear fan assembly

22 Counter-rotating compressor section (CR) 24 Counter-rotating LP compressor

26 Counter-rotating HP compressor

28 Combustion section

30 Combustor

32 Counter-rotating turbine section

34 HP counter-rotating turbine

36 Counter-rotating LP turbine

38 Unloading section

40 Fan blades

42 Fan housing

44 Rotor discs

46 Front fan shaft

47 Rear fan shaft

48 Center of gas turbine engine

50 External housing

52 Rotor or central shaft

54 External rotor

56 First set of aerodynamic surfaces 58 Internal rotor

60 Second set of aerodynamic surfaces 62 Turbine supports

63 Stator

65 Pressurized air

66 Gas

67 Purge air

98 Crankshaft

100-500 Rotating blade assembly

102-502 Blade assembly

104-504 Disc

106-506 Fastener assembly 108-508 Internal platform

110 External platform

112-512 Shovels

114-514 Dovetail

116-516 Front edge

118-418 Rear edge

119-519 Root

120 Tip

122, 222 First band

124, 224 Second band

126-426 Headquarters

128-428 First through hole

130-430 Tubular element or bushing 131-431 Fingers

132-432 Pin

134-434 Fastener

136-436 Second through hole

138-438 End? distal

140-440 Shoulder

142-442 Overhang

144 First rib

145-245 Overlap junction 146 Second rib

148 Ends? distal

248-448 Termination

150-450 Cutout

251, 351 Wall

254 Queue

256 Taken

458 Retaining element plate 460 Through hole

562 First contact region

Claims (10)

1. A rotor blade assembly for a turbine engine having a drive shaft, the rotor blade assembly including: a disk operatively coupled to the drive shaft and comprising a seat having at least a portion of a first through hole; at least one blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; And a retainer assembly securing the disc to the at least one blade assembly, the retainer assembly including: a hollow tubular element defining a second through hole, which is aligned with the first through hole; And a pin extending through at least a portion of the second through hole and the first through hole and abutting at least a portion of the hollow tubular element. 2. The rotary vane assembly of claim 1 wherein the hollow tubular member is an expandable bushing. 3. The rotary blade assembly of claim 1, wherein the pin terminates at an end? distal and the rotary blade also includes a fastener secured to the end? distal and abutting the disc. 4. The rotary vane assembly of claim 3 wherein the fastener is one of a nut, a hydraulic fastener, a magnetic fastener, a weld, an adhesive, or an electrical connection. 5. The rotary vane assembly of claim 4 wherein the fastener is the nut and stud includes a threaded section, where the nut is threaded onto the threaded section of the pin. 6. The rotary vane assembly of claim 1 wherein the dovetail defines at least another portion of the first through hole. 7. The rotary vane assembly of claim 1 wherein the disc further includes a first band and a second band which together define the seat. 8. The rotary vane assembly of claim 7 wherein the first band and the second band are coupled at an overlap joint. 9. The rotary blade assembly of claim 7, wherein the first rib includes a first rib and the second rib includes a second rib, both the first rib and the second rib abutting a corresponding portion of the dovetail. 10. Gas turbine engine, including: an engine center defining an engine centerline and including a drive shaft and a first rotor; And a rotating blade assembly, including: a disk operatively coupled to the drive shaft and comprising a seat having at least a portion of a first through hole; at least one blade assembly having an upper platform operatively coupled to the first rotor, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; And at least one retainer assembly securing the disc to the at least one blade assembly, the at least one retainer assembly including: a hollow tubular element defining a second through hole, which is aligned with the first through hole; And a pin extending through at least a portion of the second through hole and the first through hole and abutting at least a portion of the hollow tubular element.